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1
Lu, Weiyu, Guoping Huang, Xin Xiang, Jinchun Wang, and Yuxuan Yang. "Thermodynamic and Aerodynamic Analysis of an Air-Driven Fan System in Low-Cost High-Bypass-Ratio Turbofan Engine." Energies 12, no. 10 (2019): 1917. http://dx.doi.org/10.3390/en12101917.
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In some cases, the improvement of the bypass ratio (BPR) of turbofans is pursued for military or civilian purposes owing to economic, environmental, and performance reasons, among others. However, high-BPR turbofans suffer from incompatibility of spool speed, complex structure for manufacture, development difficulty, and substantially increasing costs, especially for those with small batch production. To deal with the issues, a novel low-cost concept of high-BPR turbofan with air-driven fan (ADTF) is presented in this research. First, the problems faced by high-BPR turbofans are discussed, and the difficulties of geared turbofan (GTF), which is developed as a solution to the problems, are analyzed. A novel turbofan with potential advantages is proposed, and its basic theory is interpreted. Second, high-BPR ADTF is analyzed at the top level, and the design principle and important primary parameters are discussed. Some important concepts and criteria are proposed, enabling the comparison between ADTF and GTF. Finally, an air-driven fan system, the core part of ADTF, is exploratorily designed, and numerical simulation is performed to demonstrate its feasibility.
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Mazzawy, Robert S. "Next Generation of Transport Engines." Mechanical Engineering 132, no. 12 (2010): 54. http://dx.doi.org/10.1115/1.2010-dec-6.
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This article discusses the features of very high bypass ratio turbofans and open rotor engines. Each of these engine options has its pros and cons to consider. The very large bypass ratio turbofan engine maintains that the proven capability of containment of blade failures is inherently quieter due to ability to incorporate acoustic treatment in the fan duct and is not subject to high fan tip losses associated with direct exposure to higher cruise level flight speeds. The duct does not come for free, however, and installed weight becomes a primary concern as the increased bypass ratio drives up the engine diameter. Additionally, the fan is subject to higher local airfoil incidence when the fan nozzle un-chokes at low flight speed. The open rotor engine can achieve potentially greater improvements in propulsive efficiency than a turbofan but lacks the containment and noise reduction benefits of a duct. The rotor is also exposed to flight speed, driving up tip losses at today's accepted cruise flight speeds.
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Jakubowski, Robert. "Study of Bypass Ratio Increasing Possibility for Turbofan Engine and Turbofan with Inter Turbine Burner." Journal of KONES 26, no. 2 (2019): 61–68. http://dx.doi.org/10.2478/kones-2019-0033.
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Abstract Current trends in the high bypass ratio turbofan engines development are discussed in the beginning of the paper. Based on this, the state of the art in the contemporary turbofan engines is presented and their change in the last decade is briefly summarized. The main scope of the work is the bypass ratio growth analysis. It is discussed for classical turbofan engine scheme. The next step is presentation of reach this goal by application of an additional combustor located between high and low pressure turbines. The numerical model for fast analysis of bypass ratio grows for both engine kinds are presented. Based on it, the numerical simulation of bypass engine increasing is studied. The assumption to carry out this study is a common core engine. For classical turbofan engine bypass ratio grow is compensated by fan pressure ratio reduction. For inter turbine burner turbofan, bypass grown is compensated by additional energy input into the additional combustor. Presented results are plotted and discussed. The main conclusion is drawing that energy input in to the turbofan aero engine should grow when bypass ratio is growing otherwise the energy should be saved by other engine elements (here fan pressure ratio is decreasing). Presented solution of additional energy input in inter turbine burner allow to eliminate this problem. In studied aspect, this solution not allows to improve engine performance. Specific thrust of such engine grows with bypass ratio rise – this is positive, but specific fuel consumption rise too. Classical turbofan reaches lower specific thrust for higher bypass ratio but its specific fuel consumption is lower too. Specific fuel consumption decreasing is one of the goal set for future aero-engines improvements.
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Ingraldi, A. M., T. T. Kariya, R. J. Re, and O. C. Pendergraft. "Interference Effects of Very High Bypass Ratio Nacelle Installations on a Low-Wing Transport." Journal of Engineering for Gas Turbines and Power 114, no. 4 (1992): 809–15. http://dx.doi.org/10.1115/1.2906661.
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A twin-engine, low-wing transport model, with a supercritical wing designed for a cruise Mach number of0.77 and a lift coefficient of 0.55, was tested in the 16-Foot Transonic Tunnel at NASA Langley Research Center. The purpose of this test was to compare the wing/nacelle interference effects of superfans (very high bypass ratio turbofans, BPR ≈ 18) with the interference effects of advanced turbofans (BPR ≈ 6). Flow-through nacelles were used in this study. Forces and moments on the complete model were measured using a strain gage balance and extensive surface static pressure measuements (383 orifice locations) were made on the model’s wing, nacelles, and pylons. Data were taken at Mach numbers from 0.50 to 0.80 and model angle-of-attack was varied from −4 to +8 deg. Results of the investigation indicate that superfan nacelles can be installed with approximately the same drag penalty as conventional turbofan nacelles.
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Turan, Onder, Hakan Aydın, T. Hikmet Karakoc, and Adnan Midilli. "First Law Approach of a Low Bypass Turbofan Engine." Journal of Automation and Control Engineering 2, no. 1 (2014): 62–66. http://dx.doi.org/10.12720/joace.2.1.62-66.
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Cilgin, Mehmet Emin, and Onder Turan. "Entropy Generation Calculation of a Turbofan Engine: A Case of CFM56-7B." International Journal of Turbo & Jet-Engines 35, no. 3 (2018): 217–27. http://dx.doi.org/10.1515/tjj-2017-0053.
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Abstract Entropy generation and energy efficiency of turbofan engines become greater concern in recent years caused by rises fuel costs and as well as environmental impact of aviation emissions. This study describes calculation of entropy generation for a two-spool CFM56-7B high-bypass turbofan widely used on short to medium range, narrow body aircrafts. Entropy generation and power analyses are performed for five main engine components obtaining temperature-entropy, entropy-enthalpy, pressure-volume diagrams at ≈121 kN take-off thrust force. In the study, maximum entropy production is determined to be 0.8504 kJ/kg K at the combustor, while minimum entropy generation is observed at the low pressure compressor component with the value of 0.0025 kJ/kg K. Besides, overall efficiency of the turbofan is determined to be 14 %, while propulsive and thermal efficiencies of the engine are 35 % and 40 %, respectively. As a conclusion, this study aims to show increase of entropy due to irreversibilities and produced power dimension in engine components for commercial turbofans and aero-derivative cogeneration power plants.
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Verma, Vishwas, Gursharanjit Singh, and AM Pradeep. "The effect of inlet distortion on low bypass ratio turbofan engines." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 234, no. 8 (2020): 1395–413. http://dx.doi.org/10.1177/0954410020909190.
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Inlet flow non-uniformity, commonly known as inflow distortion, has been a long-standing problem in the history of gas turbine engines. Distortion can be present in the form of total pressure, total temperature or inflow incidence or any combinations of these. The search for better and robust performance requires engines that can sustain a large amount of inlet distortion without considerable loss in the thrust. In the present paper, the effect of total pressure distortion on a single-stage compressor and low bypass ratio fans are studied. Distortion near hub and tip in the form of step radial total pressure profiles is imposed at far upstream of the rotor leading edge. A systematic approach to qualitatively predict the performance maps in the presence of these distortions is discussed. Further, two extents of total pressure distortion are explored for constant inlet distortion intensity. Hub distortion is found to increase the stability margin, whereas tip distortion reduces it. On extending the distortion extent, hub distortion drastically reduces the stability margin, whereas a comparatively lower reduction in stability margin with tip distortion is observed. The critical distortion limit is observed by varying the inlet distortion extent. Also, it is found that downstream ducts in the bypass axial fan do not interact with the upstream fan. This can be exploited to perform independent simulations of the core engine from low bypass ratio fans. Hub distortion is found to drastically affect the duct performance owing to the presence of thicker upstream inlet boundary layer.
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Asundi, Sharanabasaweshwara A., and Syed Firasat Ali. "Parametric Study of a Turbofan Engine with an Auxiliary High-Pressure Bypass." International Journal of Turbomachinery, Propulsion and Power 4, no. 1 (2019): 2. http://dx.doi.org/10.3390/ijtpp4010002.
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A parametric study of a novel turbofan engine with an auxiliary high-pressure bypass (AHPB) is presented. The underlying motivation for the study was to introduce and explore a configuration of a turbofan engine which could facilitate clean secondary burning of fuel at a higher temperature than conventionally realized. The study was also motivated by the developments in engineering materials for high-temperature applications and the potential utility of these developments. The parametric study is presented in two phases. Phase I presents a schematic of the turbofan engine with AHPB and the mathematics of the performance parameters at various stations. The proposed engine is hypothesized to consist of three streams—core stream, low-pressure bypass (LPB) stream, and the AHPB or, simply, the high-pressure bypass (HPB) stream. Phase II delves into the performance simulation and the analysis of the results in an ideal set-up. The simulation and results are presented for performance analysis when (i) maximizing engine thrust while varying the LPB and AHPB ratios, and (ii) varying the AHPB ratio while maintaining the LPB ratio constant. The results demonstrate the variations in performance of the engine and a basis for examining its potential utility for practical applications.
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Mishra, R. K., and S. K. Jha. "Thermal Fatigue Failure of Low-Pressure Turbine Blade in a Low-Bypass Turbofan Engine." Journal of Failure Analysis and Prevention 19, no. 2 (2019): 301–7. http://dx.doi.org/10.1007/s11668-019-00622-0.
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Huang, Guoping, Xin Xiang, Chen Xia, Weiyu Lu, and Lei Li. "Feasible Concept of an Air-Driven Fan with a Tip Turbine for a High-Bypass Propulsion System." Energies 11, no. 12 (2018): 3350. http://dx.doi.org/10.3390/en11123350.
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The reduction in specific fuel consumption (SFC) is crucial for small/mid-size cost-controllable aircraft, which is very conducive to reducing cost and carbon dioxide emissions. To decrease the SFC, increasing the bypass ratio (BPR) is an important way. Conventional high-BPR engines have several limitations, especially the conflicting spool-speed requirements of a fan and a low-pressure turbine. This research proposes an air-driven fan with a tip turbine (ADFTT) as a potential device for a high-bypass propulsion system. Moreover, a possible application of this ADFTT is introduced. Thermodynamic analysis results show that an ADFTT can improve thrust from a prototype turbofan. As a demonstration, we selected a typical small-thrust turbofan as the prototype and applied the ADFTT concept to improve this model. Three-dimensional flow fields were numerically simulated through a Reynolds averaged Navier-Stokes (RANS)-based computational fluid dynamics (CFD) method. The performance of this ADFTT has the possibility of amplifying the BPR more than four times and increasing the thrust by approximately 84% in comparison with the prototype turbofan.
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Mishra, R. K., Thomas Johney, K. Srinivasan, Nandi Vaishakhi, and Bhat Raghavendra. "Failure Analysis of HP Turbine Blades in a Low Bypass Turbofan Engine." Journal of Failure Analysis and Prevention 13, no. 3 (2013): 274–81. http://dx.doi.org/10.1007/s11668-013-9674-5.
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Mishra, R. K., K. Srinivasan, Johney Thomas, Nandi Vaishakhi, and Raghavendra R. Bhat. "Investigation of LP Turbine Blade Failure in a Low Bypass Turbofan Engine." Journal of Failure Analysis and Prevention 14, no. 2 (2014): 160–66. http://dx.doi.org/10.1007/s11668-014-9793-7.
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Mishra, R. K., S. Kishorekumar, and Sunil Chandel. "Investigation of Flame Blow-Out in a Low Bypass Military Turbofan Engine." Journal of Failure Analysis and Prevention 15, no. 2 (2015): 227–32. http://dx.doi.org/10.1007/s11668-015-9946-3.
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Giesecke, Daniel, Marcel Lehmler, Jens Friedrichs, Jason Blinstrub, Lothar Bertsch, and Wolfgang Heinze. "Evaluation of ultra-high bypass ratio engines for an over-wing aircraft configuration." Journal of the Global Power and Propulsion Society 2 (October 17, 2018): 493–515. http://dx.doi.org/10.22261/jgpps.8shp7k.
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Today, main hub airports are already at their capacity limit and hence, smaller airports have become more interesting for providing point-to-point connections. Unfortunately, the use of regional airports induces an increased environmental footprint for the population living around it. In an attempt to solve the related problems, the research project Coordinated Research Centre 880 aims to examine the fundamentals of a single-aisle aircraft with active high-lift configuration powered by two geared ultra-high bypass turbofan engines mounted over the wing. Low direct operating costs, noise shielding due to the over-wing configuration, and short runway lengths are the main advantages. Highlighting the performance, economical and noise benefits of a geared ultra-high bypass engine is the key aim of this paper. This assessment includes a correspondingly adjusted aircraft. Open literature values are applied to design the two investigated bypass ratios; a reference engine with a bypass ratio of 5 and 17 respectively. This study shows that a careful selection of engine mass flow, turbine entry temperature and overall pressure ratio determines the desirable bypass ratio. The aircraft direct operating costs drop by 5.7% when comparing the designed conventional with a future ultra-high bypass ratio engine. Furthermore, the sound at source for a selected mission and operating condition can be reduced by 7 dB. A variable bypass nozzle area for the ultra-high bypass ratio engine is analysed in terms of performance and operability. An increase of safety margin is shown for the turbofan engine with a variable bypass nozzle. It is concluded that this unconventional aircraft configuration with ultra-high bypass ratio engines mounted over the wing has the potential to relieve main hub airports and reduce the environmental impact.
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Choi, Won, Il-Woo Lee, and Jun-Ho Yang. "The Performance Modeling of a Low Bypass Turbofan Engine with Afterburner for Supersonic Aircraft." Journal of the Korean Society for Aeronautical & Space Sciences 39, no. 3 (2011): 269–78. http://dx.doi.org/10.5139/jksas.2010.39.3.269.
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Turan, Onder, and Hakan Aydin. "Exergy-based Sustainability Analysis of a Low-bypass Turbofan Engine: A Case Study for JT8D." Energy Procedia 95 (September 2016): 499–506. http://dx.doi.org/10.1016/j.egypro.2016.09.075.
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Stuermer, Arne. "DLR TAU-Code uRANS Turbofan Modeling for Aircraft Aerodynamics Investigations." Aerospace 6, no. 11 (2019): 121. http://dx.doi.org/10.3390/aerospace6110121.
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In the context of an increased focus on fuel efficiency and environmental impact, turbofan engine developments continue towards larger bypass ratio engine designs, with Ultra-High Bypass Ratio (UHBR) engines becoming a likely power plant option for future commercial transport aircraft. These engines promise low specific fuel consumption at the engine level, but the resulting size of the nacelle poses challenges in terms of the installation on the airframe. Thus, their integration on an aircraft requires careful consideration of complex engine–airframe interactions impacting performance, aeroelastics and aeroacoustics on both the airframe and the engine sides. As a partner in the EU funded Clean Sky 2 project ASPIRE, the DLR Institute of Aerodynamics and Flow Technology is contributing to an investigation of numerical analysis approaches, which draws on a generic representative UHBR engine configuration specifically designed in the frame of the project. In the present paper, project results are discussed, which aimed at demonstrating the suitability and accuracy of an unsteady RANS-based engine modeling approach in the context of external aerodynamics focused CFD simulations with the DLR TAU-Code. For this high-fidelity approach with a geometrically fully modeled fan stage, an in-depth study on spatial and temporal resolution requirements was performed, and the results were compared with simpler methods using classical engine boundary conditions. The primary aim is to identify the capabilities and shortcomings of these modeling approaches, and to develop a best-practice for the uRANS simulations as well as determine the best application scenarios.
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Qiu, S., H. Liu, and WP Li. "Turbofan duct geometry optimization for low noise using remote continuous adjoint method." Proceedings of the Institution of Mechanical Engineers, Part C: Journal of Mechanical Engineering Science 229, no. 1 (2014): 69–90. http://dx.doi.org/10.1177/0954406214532631.
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In this paper, a remote continuous adjoint-based acoustic propagation (RABAP) method is proposed for low noise turbofan duct design. The goal is to develop a set of adjoint equations and their corresponding boundary conditions in order to quantify the influence of geometry modifications on the amplitude of sound pressure at a near-field location. The governing equations for the 2.5D acoustic perturbation equation solver (APE) formulation for duct acoustic propagation is first introduced. This is followed by the formulation and discretization of the remote continuous adjoint equations based on 2.5D APE. The special treatment of the adjoint boundary condition to obtain sensitivities derivatives is also discussed. The theory is applied to acoustic design of an axisymmetric fan bypass duct for two different tone noise radiations. The 2.5D APE is further validated using comparisons to an experiment data of the TURNEX nozzle geometry. The implementation of the remote continuous adjoint method is validated by comparing the sensitivity derivative with that obtained using finite difference method. The result obtained confirms the effectiveness and efficiency of the proposed RABAP framework.
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Qiu, S., WB Song, and H. Liu. "Shape optimization of a general bypass duct for tone noise reduction using continuous adjoint method." Proceedings of the Institution of Mechanical Engineers, Part C: Journal of Mechanical Engineering Science 228, no. 1 (2013): 119–34. http://dx.doi.org/10.1177/0954406213481915.
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A novel continuous adjoint-based acoustic propagation method is proposed for low-noise turbofan duct design. A fan bypass duct tonal noise propagation model that is verified by comparison with an analytical solution of the modal radiation from a semi-infinite duct with the shear layer is enhanced with its continuous adjoint formulation, having been applied to design the bypass duct. First, this article presents the complete formulation of the time-dependent optimal design problem. Second, a continuous adjoint-based acoustic propagation method for two-dimensional bypass duct configurations is derived and presented. This article aims at describing the potential of the adjoint technique for aeroacoustic shape optimization. The implementation of the unsteady aeroacoustic adjoint method is validated by comparing the sensitivity derivative with that obtained by finite differences. Using a continuous adjoint formulation, the necessary aerodynamic gradient information is obtained with large computational savings over traditional finite-difference methods. The examples presented demonstrate that the combination of a continuous-adjoint algorithm with a noise prediction method can be an efficient design tool in the bypass duct noise design problem.
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Hall, Cesare A., and Daniel Crichton. "Engine Design Studies for a Silent Aircraft." Journal of Turbomachinery 129, no. 3 (2006): 479–87. http://dx.doi.org/10.1115/1.2472398.
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The Silent Aircraft Initiative is a research project funded by the Cambridge-MIT Institute aimed at reducing aircraft noise to the point where it is imperceptible in the urban environments around airports. The propulsion system being developed for this project has a thermodynamic cycle based on an ultrahigh bypass ratio turbofan combined with a variable area exhaust nozzle and an embedded installation. This cycle has been matched to the flight mission and thrust requirements of an all-lifting body airframe, and through precise scheduling of the variable exhaust nozzle, the engine operating conditions have been optimized for maximum thrust at top-of-climb, minimum fuel consumption during cruise, and minimum jet noise at low altitude. This paper proposes engine mechanical arrangements that can meet the cycle requirements and, when installed in an appropriate airframe, will be quiet relative to current turbofans. To reduce the engine weight, a system with a gearbox, or some other form of shaft speed reduction device, is proposed. This is combined with a low-speed fan and a turbine with high gap-chord spacing to further reduce turbomachinery source noise. An engine configuration with three fans driven by a single core is also presented, and this is expected to have further weight, fuel burn, and noise benefits.
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Chen, Yuwen, Zhenggui Zhou, and Cui Cui. "Aerodynamic design of an ultra-highly loaded booster of a high bypass ratio turbofan." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 233, no. 1 (2017): 240–49. http://dx.doi.org/10.1177/0954410017728974.
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Based on an existing high bypass ratio turbofan, this paper uses a new concept of diffusion blade profiles with large camber and low flow losses to design a rotor for the booster, which characterizes the low wheel speed, thus achieving ultra-high loading of the rotor, effectively increasing its stage pressure ratio and reducing the weight of the booster. A booster with only one stage is designed by using the new blade profiles to replace the original booster with three stages. A two-dimensional design method of S1 (blade to blade)/S2 (hub to tip) stream surfaces is applied to design the blades of the rotor and the stator in the booster, and the two-dimensional blade profiles in the S1 stream surface are designed by means of an optimization design method. The flow fields in the original and newly designed boosters are simulated by using a numerical method. At the design point, the newly designed booster has nearly the same total pressure ratio and mass flow rate as the original booster and exhibits higher efficiency than the original booster. The surge margins of both boosters are nearly equal, and the newly designed booster can preferably match the original intermediate case.
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Lee, Incheol, Yingzhe Zhang, and Dakai Lin. "Empirical estimation of engine-integration noise for high bypass ra-tio turbofan engines." INTER-NOISE and NOISE-CON Congress and Conference Proceedings 263, no. 2 (2021): 4511–19. http://dx.doi.org/10.3397/in-2021-2723.
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To investigate the impact of installation on jet noise from modern high-bypass-ratio turbofan engines, a model-scale noise experiment with a jet propulsion system and a fuselage model in scale was conducted in the anechoic wind tunnel of ONERA, CEPRA 19. Two area ratios (an area of the secondary nozzle over an area of the primary nozzle), 5 and 7, and various airframe configurations such as wing positions relative to the tip of the engine nacelle and flap angles, were considered. Based on the analysis of experimental data, an empirical model for the prediction of engine installation noise was proposed. The model comprises two components: one is the interaction be-tween the jet and the pressure side of the wing, and the other is the interaction between the jet and the flap tip. The interaction between the jet and the pressure side of the wing contributes to the noise at the low frequencies (≤ 1.5 kHz), and the interaction between the jet and the flap tip con-tributes to the noise at the high frequencies. The proposed model showed a good agreement with the experimental data.
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Dunn, M. G., R. M. Adams, and V. S. Oxford. "Response of Large Turbofan and Turbojet Engines to a Short-Duration Overpressure." Journal of Engineering for Gas Turbines and Power 111, no. 4 (1989): 740–47. http://dx.doi.org/10.1115/1.3240321.
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The influences of thrust setting and overpressure level on engine operating characteristics have been obtained for two different high-thrust engines. The thrust setting was varied from engine-off to take-off rated thrust (TRT) and the overpressure was varied from 6.9 kPa (1.0 psi) to 19.4 kPa (2.8 psi). The specific engines under consideration were the Pratt/Whitney TF33 low bypass ratio turbofan and the Pratt/Whitney J57 turbojet. The experimental results suggest that overpressure has little influence on either the HP compressor speed or the exhaust gas total temperature. However, the magnitude of the overpressure has a large influence on turbine exhaust total pressure and on the inlet casing and the diffuser casing radial displacements. The J57 turbine casing was significantly influenced by the overpressure, whereas the TF33 turbine casing was relatively insensitive. The J57 inlet casing radial displacement was noticeably greater than the corresponding turbofan displacement. Even though the component radial displacements for the TF33 exceeded the steady-state red-line limit by more than 300 percent, the engine did not sustain any permanent damage. The J57 did, however, experience an internal rub at an overpressure of about 14.5 kPa (2.1 psi).
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Mark, C. Priyant, and A. Selwyn. "Design and analysis of annular combustion chamber of a low bypass turbofan engine in a jet trainer aircraft." Propulsion and Power Research 5, no. 2 (2016): 97–107. http://dx.doi.org/10.1016/j.jppr.2016.04.001.
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Acarer, Sercan, and Ünver Özkol. "An extension of the streamline curvature through-flow design method for bypass fans of turbofan engines." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 231, no. 2 (2016): 240–53. http://dx.doi.org/10.1177/0954410016636159.
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The two-dimensional through-flow modeling of turbomachinery is still one of the most powerful tools available to the turbomachinery industry for aerodynamic design, analysis, and post-processing of test data due to its robustness and speed. Although variety of aspects of such a modeling approach are discussed in the publicly available literature for compressors and turbines, not much emphasis is placed on combined modeling of the fan and the downstream splitter of turbofan engines. The current article addresses this void by presenting a streamline curvature through-flow methodology that is suitable for inverse design for such a problem. A new split-flow method for the streamline solver, alternative to the publicly available analysis-oriented method, is implemented and initially compared with two-dimensional axisymmetric computational fluid dynamics on two representative geometries for high and low bypass ratios. The empirical models for incidence, deviation, loss, and end-wall blockage are compiled from the literature and calibrated against two test cases: experimental data of NASA two-stage fan and three-dimensional computational fluid dynamics of a custom-designed transonic fan stage. Finally, experimental validation against GE-NASA bypass fan case is accomplished to validate the complete methodology. The proposed method is a simple extension of streamline curvature method and can be applied to existing compressor methodologies with minimum numerical effort.
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RICHTER, CHRISTOPH, HANNES LÜCK, ŁUKASZ PANEK, and FRANK THIELE. "METHODS FOR SUPPRESSING SHEAR LAYER INSTABILITIES FOR CAA." Journal of Computational Acoustics 19, no. 02 (2011): 181–203. http://dx.doi.org/10.1142/s0218396x11004420.
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The rearward propagation of tonal noise from the main fan and the engine core of modern high bypass aeroengines is one of the current demanding applications of CAA methods. One of the main features of this problem is the radiation of tones from main fan and turbine through the shear layers of core and bypass jets. This can approximately be described by a solution of the linearized perturbed Euler equations over a sheared turbulent averaged base flow field. However, these equations not only describe sound propagation, but also provide a stability analysis for the sheared base flow. Three techniques with the potential to calculate an acoustic solution and at the same time to suppress the instability are compared in this paper. The radiation of a source from a two-dimensional hot jet, chosen from a CAA workshop on benchmark problems, is considered first. Then, the techniques are adopted for the simulation of a single azimuthal mode radiating from the bypass duct of a turbofan engine, as an example for the realistic application. The first technique is based on filtering the mean flow field, over which the perturbation equations are solved. A low-order filter is applied. Subsequently, an adaption of this method, which considers a filtering of the mean flow derivatives in addition, is proved to be very beneficial. The result then reflects the analytical solution of the benchmark problem very well. The second technique filters the source terms in the governing equations. In a first attempt, all mean flow derivatives are neglected to suppress the instability. A more physical motivated variant of the approach neglects only source terms in the momentum equations. However, both provide unsatisfactory predictions of the acoustic field for the benchmark. Finally, a third technique is implemented, which considers the modification of the velocity derivatives in the momentum equations, as this method has demonstrated one of the best predictions for the benchmark problem. Nevertheless, the latter technique has no 3D extension and thus fails in suppressing the instability waves in the turbofan application.
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Mund, Friederike C., Georgios Doulgeris, and Pericles Pilidis. "Enhanced Gas Turbine Performance Simulation Using CFD Modules in a 2D Representation of the Low-Pressure System for a High-Bypass Turbofan." Journal of Engineering for Gas Turbines and Power 129, no. 3 (2006): 761–68. http://dx.doi.org/10.1115/1.2364197.
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The improvement of performance simulation for gas turbines has been approached in very different ways. In particular for high-bypass turbofans, efforts have been made to investigate radial flow distributions. The aim of the presented study was to combine a conventional characteristics-based performance code using a 2D representation of the fan with 2D representations of the adjoining intake and bypass system. Computational fluid dynamics (CFD) was the chosen tool to generate modules for the intake, bypass duct, and bypass nozzle. This approach required geometry data. A design procedure to generate these components in an axisymmetric meridional fashion and the numerical requirements for the CFD modules were developed. Typical component performances were predicted and the combined use of CFD and the performance code showed that in terms of performance, the inclusion of intake and bypass losses and the radial inlet distribution was worth considering. In particular, however, the required numerical effort was significant.
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Liang, Zhi Chao, Jie Hong, Yan Hong Ma, and Tian Yuan He. "FEM Modeling Technology and Vibration Analysis of Flexible Rotor System." Applied Mechanics and Materials 226-228 (November 2012): 257–61. http://dx.doi.org/10.4028/www.scientific.net/amm.226-228.257.
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The low-pressure spool in a high-bypass ratio turbofan engine has its unique characteristics. The slightness rotor and the large lumped mass in blade and disk are the structural characteristics while the dynamic coupling between bladed-disk and shaft is the mechanical characteristic. This paper studies on the modeling technology of a flexible rotor system based on the finite element method (FEM). The equivalent-disk method, an equivalent principle of finite element modeling, is put forward. The blades are equivalent to a disk which can not only retain the structural and mechanical characteristics but also control the scale of the model and convert the periodic geometry into axisymmetric geometry. The results show that the equivalent-disk method is helpful to improve the modeling techniques of rotor system and can be used in engineering.
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Shao, Wan Ren, Xi Hai Xu, Jing Yu He, and Fei Wu. "Study of Jet Noise Reduction on Separated Exhaust System Using Chevron Nozzles." Advanced Materials Research 1078 (December 2014): 228–34. http://dx.doi.org/10.4028/www.scientific.net/amr.1078.228.
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The jet noise reduction of chevron nozzles was investigated on high bypass ratio turbofan engine separated exhaust system using both computational predictions and scale model experiments. Six different exhaust nozzles are designed including one baseline nozzle and five different chevron nozzles. The jet noise experiments were carried out in the anechoic chamber. Tam and Auriault’s jet noise prediction theory and MGBK theory were used to predict the noise spectra of different exhaust nozzles. The results show that the far-field noise spectra as well as the noise reduction benefits of chevrons are predicted correctly by the two theories although some discrepancies occur at the high frequency range, and Tam and Auriault’s jet noise theory can give relatively more accurate prediction results. chevron nozzles reduce jet noise at the low frequencies, but increase it at high frequencies.
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Liu, Hongrui, Jun Liu, Qiang Du, Guang Liu, and Pei Wang. "Unsteady flow mechanism of the integrated aggressive inter-turbine duct in low Reynolds number condition." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 234, no. 9 (2020): 1507–17. http://dx.doi.org/10.1177/0954410020914786.
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Aggressive inter-turbine duct, which has ultra-high bypass ratio and ultra-short axial length, is widely applied in the modern turbofan engine because it can reduce engine weight and improve low-pressure rotor dynamic characteristics. However, the aggressive inter-turbine duct that has swirling flow, wake, shock, and tip clearance leakage flow of upstream high-pressure turbine, and even has structs in its flow channel, is liable to separate, especially in high-altitude low Reynolds number (Re) condition. In addition, its downstream low-pressure turbine is on the edge of separation too. In this paper, an integrated aggressive inter-turbine duct embedded with wide-chord low-pressure turbine nozzle is adopted to eliminate the aggressive inter-turbine duct's end-wall separation. Since there are many studies on suppressing the blade suction surface's separation by upstream wake, in this study inherent wake is utilized to suppress the boundary layer separation on low-pressure turbine nozzle's suction surface in the integrated aggressive inter-turbine duct. The paper studies the unsteady flow mechanisms of the integrated aggressive inter-turbine duct (especially the separation and transition mechanisms of low-pressure turbine nozzle's suction surface boundary layer) by the computatioinal simulation method.
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Guo, Jin, Jun Hu, Baofeng Tu, and Zhiqiang Wang. "A mixed-fidelity computational model of aero engine for inlet distortion." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 233, no. 14 (2019): 5295–309. http://dx.doi.org/10.1177/0954410019841798.
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This paper presents a mixed-fidelity model to investigate the effect of complex inlet distortion on aero gas turbine engines. The approach is developed by integrating a three-dimensional body force model of multistage compressors and a classical two-dimensional parallel engine model. The internal flow field of a high bypass ratio turbofan engine under different forms of total pressure distortion is simulated by the model. The simulations are discussed in depth to reveal the features of the distorted flow field in the compression components under the whole engine environment. The relationship between the overall performance of the engine and the intensity of inlet total pressure distortion is investigated quantitatively. It is evident that the mixed-fidelity model has the capability to quantitatively evaluate the effect of complex inlet distortion on the performance and the internal flow field of engines with low computational costs.
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RIENSTRA, SJOERD W., and WALTER EVERSMAN. "A numerical comparison between the multiple-scales and finite-element solution for sound propagation in lined flow ducts." Journal of Fluid Mechanics 437 (June 22, 2001): 367–84. http://dx.doi.org/10.1017/s0022112001004438.
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An explicit, analytical, multiple-scales solution for modal sound transmission through slowly varying ducts with mean flow and acoustic lining is tested against a numerical finite-element solution solving the same potential flow equations. The test geometry taken is representative of a high-bypass turbofan aircraft engine, with typical Mach numbers of 0.5–0.7, circumferential mode numbers m of 10–40, dimensionless wavenumbers of 10–50, and both hard and acoustically treated inlet walls of impedance Z = 2 − i. Of special interest is the presence of the spinner, which incorporates a geometrical complexity which could previously only be handled by fully numerical solutions. The results for predicted power attenuation loss show in general a very good agreement. The results for iso-pressure contour plots compare quite well in the cases where scattering into many higher radial modes can occur easily (high frequency, low angular mode), and again a very good agreement in the other cases.
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Kaiser, Sascha, Markus Nickl, Christina Salpingidou, Zinon Vlahostergios, Stefan Donnerhack, and Hermann Klingels. "Investigations of the synergy of Composite Cycle and intercooled recuperation." Aeronautical Journal 122, no. 1252 (2018): 869–88. http://dx.doi.org/10.1017/aer.2018.46.
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ABSTRACTThe synergistic combination of two promising engine architectures for future aero engines is presented. The first is the Composite Cycle Engine, which introduces a piston system in the high pressure part of the core engine, to utilise closed volume combustion and high temperature capability due to instationary operation. The second is the Intercooled Recuperated engine that employs recuperators to utilise waste heat from the core engine exhaust and intercooler to improve temperature levels for recuperation and to reduce compression work. Combinations of both architectures are presented and investigated for improvement potential with respect to specific fuel consumption, engine weight and fuel burn against a turbofan. The Composite Cycle alone provides a 15.6% fuel burn reduction against a turbofan. Options for adding intercooler were screened, and a benefit of up to 1.9% fuel burn could be shown for installation in front of a piston system through a significant, efficiency-neutral weight decrease. Waste heat can be utilised by means of classic recuperation to the entire core mass flow before the combustor, or alternatively on the turbine cooling bleed or a piston engine bypass flow that is mixed again with the main flow before the combustor. As further permutation, waste heat can be recovered either after the low pressure turbine – with or without sequential combustion – or between the high pressure and low pressure turbine. Waste heat recovery after the low pressure turbine was found to be not easily feasible or tied to high fuel burn penalties due to unfavourable temperature levels, even when using sequential combustion or intercooling. Feasible temperature levels could be obtained with inter-turbine waste heat recovery but always resulted in at least 0.3% higher fuel burn compared to the non-recuperated baseline under the given assumptions. Consequently, only the application of an intercooler appears to provide a considerable benefit for the examined thermodynamic conditions in the low fidelity analyses of various engine architecture combinations with the specific heat exchanger design. Since the obtained drawbacks of some waste heat utilisation concepts are small, innovative waste heat management concepts coupled with the further extension of the design space and the inclusion of higher fidelity models may achieve a benefit and motivate future investigations.
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Nakao, M., M. Ikeyama, and S. Abe. "Analytical Condition Inspection and Extension of Time Between Overhaul of F3-30 Engine." Journal of Engineering for Gas Turbines and Power 114, no. 2 (1992): 196–200. http://dx.doi.org/10.1115/1.2906572.
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F3-30 is the low-bypass-ratio turbofan engine developed to power the T-4 intermediate trainer for the Japan Air Self Defense Force (JASDF). The actual field service was started in Sept., 1988. The program to extend time between overhaul (TBO) of the F3-30 has been running. Analytical condition inspection (ACI) and accelerated mission testing (AMT) were conducted to confirm sufficient durability to extend TBO. Most deteriorations of parts and performance due to AMT were also found by ACI after field operation with approximately the same deterioration rate. On the other hand, some deteriorations were found by ACI only. These results show that ACI after field operation is also necessary to confirm the TBO extension, although AMT simulates the deterioration in field operations very well. The deteriorations that would be caused by the field operation during one extended-TBO were estimated with the results of ACI and AMT, and it was concluded that the F3-30 has sufficient durability for TBO extension to the next step.
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Hamed, Awatef A., Widen Tabakoff, Richard B. Rivir, Kaushik Das, and Puneet Arora. "Turbine Blade Surface Deterioration by Erosion." Journal of Turbomachinery 127, no. 3 (2004): 445–52. http://dx.doi.org/10.1115/1.1860376.
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This paper presents the results of a combined experimental and computational research program to investigate turbine vane and blade material surface deterioration caused by solid particle impacts. Tests are conducted in the erosion wind tunnel for coated and uncoated blade materials at various impact conditions. Surface roughness measurements obtained prior and subsequent to the erosion tests are used to characterize the change in roughness caused by erosion. Numerical simulations for the three-dimensional flow field and particle trajectories through a low-pressure gas turbine are employed to determine the particle impact conditions with stator vanes and rotor blades using experimentally based particle restitution models. Experimental results are presented for the measured blade material/coating erosion and surface roughness. The measurements indicate that both erosion and surface roughness increase with impact angle and particle size. Computational results are presented for the particle trajectories through the first stage of a low-pressure turbine of a high bypass turbofan engine. The trajectories indicate that the particles impact the vane pressure surface and the aft part of the suction surface. The impacts reduce the particle momentum through the stator but increase it through the rotor. Vane and blade surface erosion patterns are predicted based on the computed trajectories and the experimentally measured blade coating erosion characteristics.
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Fondelli, Tommaso, Antonio Andreini, Riccardo Da Soghe, Bruno Facchini, and Lorenzo Cipolla. "Numerical Simulation of Oil Jet Lubrication for High Speed Gears." International Journal of Aerospace Engineering 2015 (2015): 1–13. http://dx.doi.org/10.1155/2015/752457.
Full textAbstract:
The Geared Turbofan technology is one of the most promising engine configurations to significantly reduce the specific fuel consumption. In this architecture, a power epicyclical gearbox is interposed between the fan and the low pressure spool. Thanks to the gearbox, fan and low pressure spool can turn at different speed, leading to higher engine bypass ratio. Therefore the gearbox efficiency becomes a key parameter for such technology. Further improvement of efficiency can be achieved developing a physical understanding of fluid dynamic losses within the transmission system. These losses are mainly related to viscous effects and they are directly connected to the lubrication method. In this work, the oil injection losses have been studied by means of CFD simulations. A numerical study of a single oil jet impinging on a single high speed gear has been carried out using the VOF method. The aim of this analysis is to evaluate the resistant torque due to the oil jet lubrication, correlating the torque data with the oil-gear interaction phases. URANS calculations have been performed using an adaptive meshing approach, as a way of significantly reducing the simulation costs. A global sensitivity analysis of adopted models has been carried out and a numerical setup has been defined.
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Eggers, Torben, Hye Rim Kim, Simon Bittner, Jens Friedrichs, and Joerg R. Seume. "Aerodynamic and Aeroelastic Effects of Design-Based Geometry Variations on a Low-Pressure Compressor." International Journal of Turbomachinery, Propulsion and Power 5, no. 4 (2020): 26. http://dx.doi.org/10.3390/ijtpp5040026.
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In modern aircraft engines, the low-pressure compressor (LPC) is subjected to a flow characterized by strong wakes and secondary flows from the upstream fan. This concerns ultra-high bypass ratio (UHBR) turbofan engines, in particular. This paper presents the aerodynamic and aeroelastic sensitivities of parametric variations on the LPC, driven by the design considerations in the upstream fan. The goal of this investigation was to determine the influence of design-based geometry parameter variations on the LPC performance under realistic inlet flow distributions and the presence of an s-duct. Aerodynamic simulations are conducted at the design and off-design operating points with the fan outflow as the inlet boundary conditions. Based on the aerodynamic results, time-linearized unsteady simulations are conducted to evaluate the vibration amplitude at the resonance operating points. First, the bypass ratio is varied by reducing the channel height of the LPC. The LPC efficiency decreases by up to 1.7% due to the increase in blockage of the core flow. The forced response amplitude of the rotor decreases with increasing bypass ratio due to increased aerodynamic damping. Secondly, the fan cavity leakage flow is considered as it directly affects the near hub fan flow and thus the inflow of the LPC. This results in an increased total-pressure loss for the s-duct due to mixing losses. The additional mixing redistributes the flow at the s-duct exit leading to a total-pressure loss reduction of 4.3% in the first rotor at design point. This effect is altered at off-design conditions. The vibration amplitude at low speed resonance points is increased by 19% for the first torsion and 26% for second bending. Thirdly, sweep and lean are applied to the inlet guide vane (IGV) upstream of the LPC. Despite the s-duct and the variable inlet guide vane (VIGV) affecting the flow, the three-dimensional blade design achieves aerodynamic and aeroelastic improvements of rotor 1 at off-design. The total-pressure loss reduces by up to 18% and the resonance amplitude more than 10%. Only negligible improvements for rotor 1 are present at the design point. In a fourth step, the influence of axial gap size between the stator and the rotor rows in the LPC is examined in the range of small variations which shows no distinct aerodynamic and aeroelastic sensitivities. This finding not only supports previous studies, but it also suggests a correlation between mode shapes and locally increased excitaion with increasing axial gap size. As a result, potential design improvements in future fan-compressor design are suggested.
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Римаренко, Євген Олександрович. "СУЧАСНИЙ СТАН ПРОБЛЕМИ РОЗРОБКИ ЗВУКОПОГЛИНАЮЧИХ КОНСТРУКЦІЙ ДЛЯ ГАЗОТУРБІННИХ ДВИГУНІВ". Aerospace technic and technology, № 8 (31 серпня 2020): 111–20. http://dx.doi.org/10.32620/aktt.2020.8.15.
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A review of current international aviation noise requirements is provided. It is shown that international requirements for aircraft noise levels are constantly increasing. The rapid growth of the international fleet leads to an increase in the number of take-offs and landings of aircraft at airports, and as a result, the problem of aircraft noise still relevant. To reduce noise levels at airports, various methods are used, one of which is the operational limitations of aircraft concerning the levels of noise they create. In European Union countries there are operational restrictions for aircraft meeting the requirements of Chapter 3 with a noise margin of less than 10 EPNdB. For already established aircraft that have passed certification, again it is necessary to look for methods to reduce noise. The main type of aircraft in operation in the world is an aircraft with turbofan engines. For such an aircraft, the main sources of noise during take-off will be the noise of the fan and jet, while landing, the noise of the landing gear, flaps, slats, and fan noise. When choosing a method of reducing aircraft noise, it should determine the source that most affects the overall noise level. It has been determined that fan noise is one of the main sources of noise. Acoustic liners constructions are widely used to reduce the noise level created by the fan. They are one of the most priority areas for reducing fan noise. Achievements in the use of acoustic liners to decrease the noise of domestic aircraft An-124-100, An-148-100 are considered. It is noted that due to the increasing requirements for aircraft noise, it is necessary to use new acoustic liners with improved sound-absorbing properties. It was determined that it is possible to improve the sound-absorbing properties of the acoustic liners by expanding the frequency range of sound absorption of such structures.Modern methods for improving the acoustic properties of the acoustic liners are presented: the use of multilayer resonant acoustic liners makes it possible to customize the design for an increased number of calculation parameters; the use of modified variants of the acoustic liners core such as corrugated core, oversized perforated core; the use of porous and porous fiber materials in the design of the acoustic liners to provide additional sound-absorbing ability, the use of low-frequency acoustic liners to reduce the noise of promising turbofan engines with a high and ultra-high bypass ratio.
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Keller, M. C., C. Kromer, L. Cordes, C. Schwitzke, and H. J. Bauer. "CFD study of oil-jet gear interaction flow phenomena in spur gears." Aeronautical Journal 124, no. 1279 (2020): 1301–17. http://dx.doi.org/10.1017/aer.2020.44.
Full textAbstract:
ABSTRACTOil-jet lubrication and cooling of high-speed gears is frequently employed in aeronautical systems, such as novel high-bypass civil aero engines based on the geared turbofan technology. Using such oil-jet system, practitioners aim to achieve high cooling rates on the flanks of the highly thermally loaded gears with minimum oil usage. Thus, for an optimal design, detailed knowledge about the flow processes is desired. These involve the oil exiting the nozzle, the oil impacting on the gear teeth, the oil spreading on the flanks, the subsequent oil fling-off, as well as the effect of the design parameters on the oil flow. Better understanding of these processes will improve the nozzle design phase, e.g. regarding the nozzle positioning and orientation, as well as the nozzle sizing and operation.Most related studies focus on the impingement depth to characterize the two-phase flow. However, the level of information of this scalar value is rather low for a complete description of the highly dynamic three-dimensional flow. Motivated by the advancements in numerical methods and the computational resources available nowadays, the investigation of the oil-jet gear interaction by means of computational fluid dynamics (CFD) has come into focus lately.In this work, a numerical setup based on the volume-of-fluid method is presented and employed to investigate the two-phase flow phenomena occurring in the vicinity of the gear teeth. The setup consists of a single oil-jet impinging on a single rotating spur gear. By introducing new metrics for characterizing the flow phenomena, extensive use of the possibilities of modern CFD is made, allowing a detailed transient and spatially resolved flow analysis. Thus, not only the impingement depth, but also the temporal and spatial evolution of wetted areas on the gear flanks, as well as the evolution of the oil volume in contact with the gear flanks are extracted from the simulation data and compared in a CFD study.The study consists of 21 different simulation cases, whereby the effect of varying the jet velocity, the jet inclination angle, the jet diameter, and the gear speed are examined. Consistent results compared to a simplified analytical approach for the impinging depth are obtained and the results for the newly introduced metrics are presented.
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Balan, C. "Design Considerations of a Versatile Simulator for High-Bypass Turbofans." Journal of Engineering for Gas Turbines and Power 117, no. 1 (1995): 31–37. http://dx.doi.org/10.1115/1.2812778.
Full textAbstract:
The continuing requirement for performance improvement of higher thrust turbofans is met by increased bypass ratios. The trend toward higher bypass ratios and relatively large-diameter low-pressure-ratio fans requires innovative design approaches, which include shorter inlets, slimmer nacelles, shorter fan ducts and exhaust systems, and possible elimination of thrust reverser. The success of this new generation of high-bypass ducted turbofans depends on understanding the acoustic impact from reduced treatment areas, inlet-fan coupling and operability, as well as overall system performance. To achieve these goals a versatile scale model propulsion simulator large enough to operate as a fan rig, yet small enough to be installed in a wind tunnel for evaluating overall acoustic, operability, and system performance, was developed. The criteria of designing such a simulator and its capabilities are discussed in this paper.
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Riegler, Claus, Michael Bauer, and Joachim Kurzke. "Some Aspects of Modeling Compressor Behavior in Gas Turbine Performance Calculations." Journal of Turbomachinery 123, no. 2 (2000): 372–78. http://dx.doi.org/10.1115/1.1368123.
Full textAbstract:
Performance calculation procedures for gas turbine engines are usually based on the performance characteristics of the engine components, and especially the turbo components are of major interest. In this paper methods of modelling compressors in gas turbine performance calculations are discussed. The basic methodologies based on Mach number similarity are summarized briefly including some second order effects. Under extreme engine partload conditions, as for example subidle or windmilling, the operating points in the compressor map are located in a region which is usually not covered by rig tests. In addition the parameters usually used in compressor maps are no longer appropriate. For these operating conditions a method is presented to extrapolate compressor maps towards very low spool speed down to the locked rotor. Instead of the efficiency more appropriate parameters as for example specific work or specific torque are suggested. A compressor map prepared with the proposed methods is presented and discussed. As another relevant topic the performance modelling of fans for low bypass ratio turbofans is covered. Due to the flow splitter downstream of such a fan the core and bypass stream may be throttled independently during engine operation and bypass ratio becomes a third independent parameter in the map. Because testing a fan on the rig for various bypass ratios is a very costly task, a simplified method has been developed which accounts for the effects of bypass ratio.
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Moreau, Stéphane, and Michel Roger. "Advanced noise modeling for future propulsion systems." International Journal of Aeroacoustics 17, no. 6-8 (2018): 576–99. http://dx.doi.org/10.1177/1475472x18789005.
Full textAbstract:
In order to meet noise specifications for future foreseen aircraft propulsion systems, such as for ultrahigh bypass ratio turbofans and contra-rotating open rotors, the dominant turbomachinery noise mechanisms need to be modeled accurately at an early design stage. Two novel methods are presented here, which could significantly improve the existing analytical noise models. For the high-solidity ultrahigh bypass ratio, a mode-matching technique based on a modal expansion of acoustic and vortical variables in each subdomain of a blade row is shown to accurately reproduce sound generation and propagation in two-dimensional bifurcated channels and in three-dimensional annular unstaggered flat-plate cascades. For the low solidity contra-rotating open rotors, several extensions to Amiet’s compressible isolated airfoil theory are coupled with Curle’s and Ffowcs Williams and Hawkings’ acoustic analogy in the frequency domain within a strip theory framework, to yield both far-field tonal and broadband noise. Including sweep in both tonal and broadband noise models is shown to significantly improve the comparison with experiments on a stationary swept airfoil in a uniform turbulent stream and on a realistic contra-rotating open rotor geometry at approach conditions.
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Akinnuli, B. O., and O. J. Oladipo. "Design of Next Generation Civil and Military Aircraft with Ultra-High Bypass Engine using Composites, Advanced Materials and Technology." Mechanical Engineering Research 8, no. 2 (2018): 48. http://dx.doi.org/10.5539/mer.v8n2p48.
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Indirect combustion noise had not been attracting research in the past, but recent indication seems to prove that it could be a threat in the future if not addressed. Means of reducing this type of noise to a low decibel value was also included. Noise is due to the ingestion of distorted atmospheric turbulence, as the two set of blades rotate in different direction. Open rotor noise is higher since the rotors are fully exposed to oncoming turbulence and lack ducting or a nacelle to attenuate the radiated sound. A thorough review on the technology that can replace conventional turbofan was carried out. It was found that none of this technology can meet up with the ACARE and NASA 2020 vision but left a gap to be filled. Because open rotor is the most proven engine that is able to satisfy this requirements, different methods are adopted and integrated to reduce open rotor noise. Attention was paid to the geometry of the blade, hub and blade length, the vorticity and interaction noise are simulated until an optimized blade was achieved. The integration problem of open rotor was addressed where the engine was located to minimize perceive noise to the payload.
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Oliveira, Fábio, and Francisco Brójo. "Parameterization of a Conventional and Regenerated UHB Turbofan." Open Engineering 5, no. 1 (2015). http://dx.doi.org/10.1515/eng-2015-0030.
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AbstractThe attempt to improve aircraft engines efficiency resulted in the evolution from turbojets to the first generation low bypass ratio turbofans. Today, high bypass ratio turbofans are the most traditional type of engine in commercial aviation. Following many years of technological developments and improvements, this type of engine has proved to be the most reliable facing the commercial aviation requirements. In search of more efficiency, the engine manufacturers tend to increase the bypass ratio leading to ultra-high bypass ratio (UHB) engines. Increased bypass ratio has clear benefits in terms of propulsion system like reducing the specific fuel consumption. This study is aimed at a parametric analysis of a UHB turbofan engine focused on short haul flights. Two cycle configurations (conventional and regenerated) were studied, and estimated values of their specific fuel consumption (TSFC) and specific thrust (Fs) were determined. Results demonstrate that the regenerated cycle may contribute towards a more economic and friendly aero engines in a higher range of bypass ratio.
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Clark, Robert, Mingxuan Shi, Jonathan Gladin, and Dimitri N. Mavris. "Design and Analysis of an Aircraft Thermal Management System Linked to a Low Bypass Ratio Turbofan Engine." Journal of Engineering for Gas Turbines and Power, August 6, 2021. http://dx.doi.org/10.1115/1.4052031.
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Abstract The design of an aircraft thermal management system (TMS) that is capable of rejecting heat loads into the bypass stream of a typical low-bypass ratio turbofan engine, or a ram-air stream, is investigated. The TMS consists of an air cycle system, similar to the typical air cycle machines used on current aircraft, both military and commercial. This system turbocharges compressor bleed air and uses heat exchangers in a ram air stream, or the engine bypass stream, to cool the engine bleed air prior to expanding it to low temperatures suitable for heat rejection. In this study, a simple low-bypass ratio afterburning turbofan engine was modeled in NPSS to provide boundary conditions to the TMS system throughout the flight envelope of a typical military fighter aircraft. Two variations of the TMS system, a ram air cooled and a bypass air cooled, were sized to handle a given demanded aircraft heat load. The ability of the sized TMS to reject the demanded aircraft load throughout several key off-design points was analyzed. It was observed that the maximum load dissipation capability of the TMS is tied to the amount of engine bleed flow, while the level of bleed flow required is set by the temperature conditions imposed by the aircraft cooling system. Notably, engine bypass stream temperatures significantly limit the thermodynamic viability of a TMS designed with bypass air as the heat sink. The results demonstrate the advantage that variable cycle engines may have for future aircraft designs.
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Speak, Trevor H., Robert J. Sellick, Vadim Kloos, and Peter Jeschke. "Dual Drive Booster for a Two-Spool Turbofan: Performance Effects and Mechanical Feasibility." Journal of Engineering for Gas Turbines and Power 138, no. 2 (2015). http://dx.doi.org/10.1115/1.4031274.
Full textAbstract:
A novel two-spool turbofan engine configuration is described which uses a booster powered by both the low an high pressure spools. Design and off-design performance analysis shows the operating characteristics of the configuration, and a mechanical feasibility study of the gearbox is presented. The trends toward ever higher engine overall pressure ratio and bypass ratio have resulted in a combination of higher pressure ratio and lower blade speed in the booster compressor of conventional two-spool turbofans. This combination gives rise to many stages in the booster and/or lower booster efficiency and also a higher degree of off-design mismatch between the core compressors. The current paper describes an engine architecture which aims to alleviate both these issues by powering the booster compressor from both low and high pressure spools through an epicyclic gear system. We have called this engine architecture the dual drive booster. The concept gives the engine designer greater flexibility to optimize component performance and work split, resulting in the potential for lower cruise specific fuel consumption and higher hot-day takeoff thrust capability than current engine configurations. The gear system is described along with the mathematical derivation of the booster rotational speed in terms of LP- and HP-spool speeds. Both the design point and off-design performance modeling have been conducted and comparison is made between a conventional turbofan and a turbofan fitted with the dual drive booster. The results show a significant enhancement in takeoff thrust due to the better speed match of the booster. The paper also describes the results of a preliminary study into the design and mechanical feasibility of the engine architecture and gear system. The presented concept is an alternative to the conventional turbofan and should be considered during the conceptual design of future aircraft engines.
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Ji, Zifei, Huiqiang Zhang, Bing Wang, and Wei He. "Comprehensive Performance Analysis of the Turbofan With a Multi-Annular Rotating Detonation Duct Burner." Journal of Engineering for Gas Turbines and Power 142, no. 2 (2020). http://dx.doi.org/10.1115/1.4045518.
Full textAbstract:
Abstract The performance analysis of mixed-exhaust turbofan engine with multi-annular rotating detonation duct burner (RDDB) is conducted for the first time, considering that the flow path of the bypass duct is ideal for a rotating detonation combustor (RDC). The configuration of the multi-annular rotating detonation combustor is constructed aiming at the advantages of a wider operation range and uniform outlet parameters over the single-annular one. Then, a parametric analysis model of the mixed-exhaust turbofan engine with a rotating detonation duct burner is developed. Thereafter, the effects of duct burner parameters on the engine performance and operating characteristics are investigated. The mixed-exhaust turbofan engine with a rotating detonation duct burner shows superior overall performance to that of one with an isobaric afterburner (ICAB) over a wide operation range. The separate-exhaust rotating detonation duct burner can hold characteristics that are higher than those of the mixed-exhaust one at lower values of fan pressure ratio, while the mixed-exhaust one corresponds to lower values of turbine inlet temperature. When the rotating detonation duct burner is “on,” the low-pressure rotor operating line moves toward the surge line on the low corrected shaft speed side but away from the surge line on the high corrected shaft speed side.
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Khalid, Syed J. "Optimizing Separate Exhaust Turbofans for Cruise Specific Fuel Consumption." Journal of Engineering for Gas Turbines and Power 139, no. 12 (2017). http://dx.doi.org/10.1115/1.4037316.
Full textAbstract:
Cruise specific fuel consumption (SFC) of turbofan engines is a key metric for increasing airline profitability and for reducing CO2 emissions. Although increasing design bypass ratio (BPR) of separate exhaust turbofan configurations improves cruise SFC, further improvements can be obtained with online control actuated variable geometry modulations of bypass nozzle throat area, core nozzle throat area, and compressor variable vanes (CVV/CVG). The scope of this paper is to show only the benefits possible, and the process used in determining those benefits, and not to suggest any particular control algorithm for searching the best combination of the control effectors. A parametric cycle study indicated that the effector modulations could increase the cruise BPR, core efficiency, transmission efficiency, propulsive efficiency, and ideal velocity ratio resulting in a cruise SFC improvement of as much as 2.6% depending upon the engine configuration. The changes in these metrics with control effector variations will be presented. Scheduling of CVV is already possible in legacy digital controls; perturbation to this schedule and modulation of nozzle areas should be explored in light of the low bandwidth requirements at steady-state cruise conditions.
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García Rosa, Nicolás, Adrien Thacker, and Guillaume Dufour. "Periodic flow structures in a turbofan fan stage in windmilling." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, August 27, 2020, 095441002094829. http://dx.doi.org/10.1177/0954410020948297.
Full textAbstract:
In a fan stage under windmilling conditions, the stator operates under negative incidence, leading to flow separation, which may present an unsteady behaviour due to rotor/stator interactions. An experimental study of the unsteady flow through the fan stage of a bypass turbofan in windmilling is proposed, using hot-wire anemometry. Windmilling conditions are reproduced in a ground engine test bed by blowing a variable mass flow through a bypass turbofan in ambient conditions. Time-averaged profiles of flow coefficient are independent of the mass flow, demonstrating the similarity of velocity triangle. Turbulence intensity profiles reveal that the high levels of turbulence production due to local shear are also independent of the inlet flow. A spectral analysis confirms that the flow is dominated by the blade passing frequency, and that the separated regions downstream of the stator amplify the fluctuations locked to the BPF without adding any new frequency. Phase-locked averaging is used to capture the periodic wakes of the rotor blades at the rotor/stator interface. A spanwise behaviour typical of flows through windmilling fans is evidenced. Through the inner sections of the fan, rotor wakes are thin and weakly turbulent, and the turbulence level remains constant through the stage. The rotor wakes thicken and become more turbulent towards the fan tip, where flow separation occurs. Downstream of the stator, maximum levels of turbulence intensity are measured in the separated flow. Large periodical zones of low velocity and high turbulence intensity are observed in the outer parts of the separated stator wake, confirming the pulsating motion of the stator flow separation, locked at the blade passing frequency. Space-time diagrams show that the flow is chorochronic, and a 2 D non-linear harmonic simulation is able to capture the main interaction modes, however, the stator incidence distribution could be affected by 3 D effects.
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Liu, Yuan, Manuj Dhingra, and J. V. R. Prasad. "Benefits of Active Compressor Stability Management on Turbofan Engine Operability." Journal of Engineering for Gas Turbines and Power 131, no. 4 (2009). http://dx.doi.org/10.1115/1.3028565.
Full textAbstract:
Active compressor stability management can play a significant role in the intelligent control of gas turbine engines. The present work utilizes a computer simulation to illustrate the potential operability benefits of compressor stability management when actively controlling a turbofan engine. The simulation, called the modular aeropropulsion system simulation (MAPSS) and developed at NASA Glenn, models the actuation, sensor, controller, and engine dynamics of a twin-spool, low-bypass turbofan engine. The stability management system is built around a previously developed stability measure called the correlation measure. The correlation measure quantifies the repeatability of the pressure signature of a compressor rotor. Earlier work has used laboratory compressor and engine rig data to develop a relationship between a compressor’s stability boundary and its correlation measure. Specifically, correlation measure threshold crossing events increase in magnitude and number as the compressor approaches the limit of stable operation. To simulate the experimentally observed behavior of these events, a stochastic model based on level-crossings of an exponentially distributed pseudorandom process has been implemented in the MAPSS environment. Three different methods of integrating active stability management within the existing engine control architecture have been explored. The results show that significant improvements in the engine emergency response can be obtained while maintaining instability-free compressor operation via any of the methods studied. Two of the active control schemes investigated utilize existing scheduler and controller parameters and require minimal additional control logic for implementation. The third method, while introducing additional logic, emphasizes the need for as well as the benefits of a more integrated stability management system.